1. Field of the Invention
The present invention relates generally to fluid reaction surfaces, and more specifically to a turbine blade with a separately formed platform.
2. Description of the Related Art including information disclosed under 37 CFR 1.97 and 1.98
In a gas turbine engine, such as an aero engine used to power an aircraft or an industrial gas turbine engine used to produce electrical power, a turbine section includes a plurality of stages of rotor blades and stator vanes to extract the energy from the hot gas flow passing through. The engine efficiency can be improved by increasing the temperature of the hot gas flow entering the turbine. However, the inlet temperature is limited to the material properties of the first stage vanes and rotor blades. To improve the efficiency, complex internal cooling circuits have also been proposed to provide impingement and film cooling to these airfoils in order to allow for a higher gas flow temperature.
Turbine blades made from a single crystal material are used in order to allow for improvements in the thermal mechanical failure (TMF), life cycle fatigue (LCF), and creep over single piece investment cast blades. Single crystal blades have a unique crystal structure that provides for higher creep resistance in the spanwise direction of the blade. However, single crystal blades that are formed with the blade platforms as a single piece have very high casting failure rates which lead to higher manufacturing costs. Also, the fillet between the airfoil portion and the platform of the blade acts as a stress concentration location. The higher resulting stress on the blade can result in shortened life.
U.S. Pat. No. 3,132,841 issued to Wilder, Jr. on May 12, 1964 and entitled COMPRESSOR BLADE AND MANUFACTURE THEREOF discloses a compressor blade made from a plastic fiber reinforced airfoil portion joined to a metallic base portion that forms the composite compressor blade. The airfoil portion includes a base with an insert that forms a dovetail shaped end opposite from the blade tip. The base is a single piece that has an outer contour of any suitable form for installation on a compressor rotor or stator body such as a conventional dovetail groove in a rotor. The internal opening of the base is shaped and dimensioned to conform to the outer contours of the blade foot and preferably the immediately adjacent part of the blade proper (see column 2, lines 28 through 42 of this patent). Apparently, the blade portion is passed through the opening in the base from the bottom end of the base. The turbine blade of the present invention has several significant structural differences to the Wilder patent that is described below.
Another prior art reference, U.S. Pat. No. 2,817,490 issued to Broffitt on Dec. 24, 1957 and entitled TURBINE BUCKET WITH INTERNAL FINS describes a turbine blade having an airfoil portion with a root portion formed of two parts, each part having an inner surface that is serrated transversely to the length of the blade and an outer surface that has dovetail grooves for insertion into a rotor disk slot. The root members are joined together by brazing or soldering to form a rigid integral turbine blade. The turbine blade of the present invention also has several significant structural differences to the Broffitt patent that is described below.
An object of the present invention is to de-couple the airfoil portion from the platform of the blade in order to reduce stress concentration.
Another object of the present invention is to produce a turbine blade made from a single crystal material in which the platform is formed from a separate piece in order to reduce casting defects.
Another object of the present invention is to provide for a single crystal turbine blade that can be secured within a standard dovetail slot of a rotor disk.